Fuel pumps usable in aircraft engines are, obviously, crucial to the safe operation of the aircraft. Such pumps are commonly centrifugal pumps with impellers operating at 25,000-35,000 rpm to raise fuel pressure from around 50 psi at the pump inlet to over 900 psi at the impeller outlet. This pressure differential between the inlet and the outlet results in significant back flow or leakage back around the impeller (between the impeller and the housing) from its high pressure outlet to its low pressure inlet.
One approach to sealing against such leaking has been to minimize the clearance between the impeller and either the housing or an intermediate sealing member fixed to the housing. Even with close manufacturing tolerances, however, it has not been possible to reduce this diametral clearance to less than about 6-10 thousandths of an inch, which minimal clearance is required in view of a number of factors, including centrifugal and thermal growth, and imprecision in establishing and maintaining the axis of rotation of the impeller in the housing. As a result of this required clearance, prior art pumps moving, for example, 200 gallons per minute have commonly leaked as much as 30-40 gallons per minute of that back to the inlet.
Still other seals which have been used in the prior art include labyrinth seals which define tortuous paths in order to attempt to minimize any leakage of fluids therethrough. Such seals are shown, for example, in U.S. Pat. No. 4,269,564 and U.S. Pat. No. 3,238,534. Such pumps still can allow 15-20% leakage such as previously described.
Leakage as described above not only reduces the capacity of the pump, but also increases the horsepower required for a desired pump output. For example, during flight and ground idle when the engines require little fuel but the pump is still operating at design speeds, the action of the pump will increase the fuel temperature, which also results in a requirement of more power for the pump (e.g., from 50 horsepower to 90 horsepower). It is, of course, desirable to minimize the horsepower requirements of any aircraft component.
The present invention is directed to overcoming one or more of the problems as set forth above.